(1) Field of the Invention
The present invention relates to a blade with reduced torsional rigidity intended for a rotor, and to a rotor equipped with such a blade. More specifically, this blade is intended for an aircraft propeller, or for a tail rotor of a rotary-wing aircraft.
(2) Description of Related Art
Traditionally, a blade extends longitudinally along its span from a first extremity intended to be secured to a rotating hub of a rotor toward a second extremity known as the “free extremity”. With regard to the rotor, it is understood that the blade extends radially from the first extremity toward the second extremity.
It should be noted that the term “longitudinal” is understood as referring to the direction of the span of the blade.
Furthermore, this blade extends transversely from a leading edge toward a trailing edge of the blade, along the chord of this blade. In particular, the blade includes an outer covering provided with a first skin in the vicinity of its extrados, for convenience referred to as the “extrados skin”, and a second skin in the vicinity of its intrados, for convenience referred to as the “intrados skin”.
Each blade of a rotor—such as a supporting rotor of a rotary-wing aircraft—produces lift during the rotary movement of this rotor, allowing the aircraft to be supported and even propelled. The lift developed by the rotor is greater or lesser, depending on the pitch angle of each rotor blade. The aerodynamic angle of incidence (i.e., the angle between the chord of the profile and the airflow) of each aerodynamic profile of the blade—for convenience referred to as the “profile”—of a section that is normal to the pitch-variation axis of the blade corresponds to a near-constant called the “wedge” of the profile at the pitch angle of the blade. This pitch-variation axis generally defines the longitudinal axis of the blade.
Under these conditions, one observes—starting at a threshold angle of incidence of a given profile (that is, of a cross-section of the blade)—a detachment of the airflows, particularly in the region of the leading edge or the trailing edge of this profile. This detachment may cause aerodynamic stalling of the blade, or even a sudden drop in its lift if this phenomenon propagates and persists in an area located between two profiles that define a critical surface along the span of the blade. Moreover, the detachment of the airflows generates vortices that cause an increase in the drag coefficient of the blade, while also causing vibrations.
To limit the detachments, one solution consists of allowing the blade to twist geometrically in relation to itself.
It should be noted that the geometric twisting of a blade can be defined by the angle formed between the chord of each profile of a section of the blade with a reference plane of the blade.
In practice, each profile of the blade can be twisted in relation to the pitch-variation axis of the blade, by an angle that is defined in relation to such a reference plane.
It is understood that for a given blade trajectory, the twisting directly affects the aerodynamic incidence of each profile. Under these conditions, the term “twist law” refers to the trend, along the span of the blade, of the twist angles obtained by construction. The twist law of a blade is immutable. This twist law is the result of a compromise that is accepted in order to satisfy the optimal functioning of the rotor in all flight regimes.
This twist law applies from the first extremity to the second extremity of the blade, along an axis known as the “twist axis”, or else along a predefined curve. Obviously, the blade can be deformed due to the effect of the forces that are applied.
Today, in order to design the rotor blades of a rotary-wing aircraft, its major structural elements—which typically consist of spars, an outer covering, and possibly ribs—are sized so as to meet the performance requirements, such as, for example, the takeoff weight of the aircraft or the resistance of the blades to centrifugal forces.
Furthermore, a filling material—which usually consists of a foam—is inserted into the free spaces between these structural elements. This filling material makes it possible to impart compression rigidity in relation to the airflow applied to the covering of the blade, and also renders the blade torsionally rigid, primarily along its span.
For example, document FR 2616409 describes a blade made of composite materials that includes a longitudinal spar along the longitudinal axis of the blade. This spar is located essentially at the center of the blade, such that the blade includes two chambers, which are located between the spar and the leading edge of the blade and the trailing edge of the blade, respectively. This spar may also consist of two elements, with an additional chamber being located between these two elements. Because each chamber may be filled with foam, this blade possesses substantial mechanical strength, even in the region of torsional rigidity about the longitudinal axis of the blade.
It should be noted that document FR 2617119 describes a blade made of composite materials that includes a structural core; an approximate aerodynamic profile that ensures the structural strength of the blade; and a trim cover that forms the desired aerodynamic profile and that consists of an intrados skin and an extrados skin. Thus, a single structural core can be associated with multiple trim covers—that is, different aerodynamic profiles—with a layer of a conforming material filling the space between the structural core and the trim cover. Thus, such a blade, which consists of different aerodynamic profiles distributed along the span of the blade, has the same structural core and therefore the same mechanical characteristics, regardless of the structural cross-sections formed from those profiles. The structural core includes at least one longitudinal spar, which may be located in the vicinity of the leading edge and/or at the center of the blade; a rigid shell; and a filling material.
Document EP 0764764 relates to a blade that includes spars that are positioned both longitudinally and transversely between an intrados skin and an extrados skin of the blade, with the spaces between the spars being filled with foam. The rigidity of the blade, as obtained through the use of these spars, allows the blade to remain functional after local impacts that affect the blade and that, in particular, destroy certain areas between these spars. Such a blade that includes spars that are positioned both longitudinally and transversely possesses very substantial rigidity, in order to be able to remain functional after such impacts.
Generally, when a blade is developed, the majority of the mechanical characteristics that are desired for the effectiveness of the blade—such as its shear stability (i.e., its ability to withstand shear forces and torque), its longitudinal rigidity in the flexion direction, and its fatigue strength or even its mass, do meet the stated expectations.
Optimization of the torsional rigidity about a longitudinal axis of a blade is a design stage that demands very close attention. Indeed, if the blade lacks torsional rigidity about this longitudinal axis, it will be too flexible, such that it will be deformed and will twist excessively due to the effect of aerodynamic forces. Thus, its aerodynamic performance will be significantly degraded. Conversely, if the blade is torsionally too rigid about this longitudinal axis, it will not be deformed and will not twist due to the effect of aerodynamic forces. However, substantial amounts of vibration may occur, as well as the detachment of airflows, particularly along the leading edge or the trailing edge of the blade, thereby causing aerodynamic stalling of the blade.
This torsional rigidity of the blade can be partially offset through active modification of the twisting of the blade—that is, while the aircraft is in flight. For example, one or more movable flaps may be added locally to the extension of the trailing edge of the blade, so as to cause this active twisting due to the effect of aerodynamic forces. Piezoelectric fibers may also be incorporated into the extrados skin and/or the intrados skin of the blade. Alternatively, the extrados and intrados skins of the blade may be made, at least locally, of anisotropic composite materials. Such blades are usually referred to as “active-twist blades”, in contrast to traditional blades, whose twist is obtained by construction and is permanently set.
Obviously, such techniques for the active twisting of a blade are complex, in terms of both their manufacture and their implementation. In order to achieve the desired active twisting, these techniques also employ units or elements that are dedicated specifically to this task—such as actuators or flaps that significantly increase the mass of the blade, which is undesirable for the rotary wing of aircraft.
Furthermore, document FR 2956856 describes a so-called “adaptive-twist blade” that includes a leaf consisting of a stack of unidirectional composite layers that are anti-symmetrical in relation to a median layer, with the said stack of unidirectional composite layers being provided with a variable number of layers along the span of the blade. First and second attachment means secure this leaf to the intrados skin and to the extrados skin, respectively, of the outer covering. Each attachment means extends into the cavity along a longitudinal direction that is parallel to the span of the blade. Then, during flight, this adaptive-twist blade can twist as a result of the twisting of the leaf due to the effect of centrifugal forces. Thus, one can speak of the blade's “adaptation to twisting”—that is, its ability to twist during flight without the use of active-twist means.
The design of a blade can also be modified in order to adjust the torsional rigidity of the blade. However, such modifications face two major difficulties.
First, this torsional rigidity cannot easily be modified without affecting the other mechanical characteristics of the blade that are essential to its proper operation. Indeed, any elementary modification of the blade intended to reduce or increase its torsional rigidity will modify other mechanical and mass-related characteristics of the blade. For example, the thicknesses of the covering of the blade, as well as the composition of the covering, can be changed. Alternatively, the filling material can be changed, in order to modify the torsional rigidity of the blade. However, this modification may affect the longitudinal rigidity of the blade. Thus, such a modification intended to adjust the torsional rigidity of the blade usually entails the development of a new blade.
Next, this torsional rigidity that is unfavorable to the proper operation of the blade is usually detected during rotation tests performed on a test bench, or even during flight tests of a prototype—that is, when the development of the blade is well advanced. Consequently, modifying the torsional rigidity of a blade has a very significant negative effect on the development of the blade, by increasing not only its development cost but also the duration of the research phase.
Therefore, the goal of the present invention is to propose a blade that overcomes the above-mentioned limitations, specifically by separating the torsional rigidity of the blade from its other mechanical characteristics, so as to obtain reduced torsional rigidity of the blade about its longitudinal axis without significantly modifying its other mechanical characteristics.
According to the invention, a blade with reduced torsional rigidity intended for a rotor includes an outer covering extending along the span of the blade of a first extremal area to a second extremal area, and also structural means.
The first extremal area is located near the first extremity of the blade, which is suitable for being secured to a rotor hub, while the second extremal area is located in the vicinity of the second, free extremity of the blade.
This blade with reduced torsional rigidity is intended particularly for an aircraft propeller blade or for a rotor of a rotary-wing aircraft, and, more specifically, for a tail rotor.
According to a first embodiment of the invention, this blade with reduced torsional rigidity is particularly well suited to an adaptive-twist blade, as described in document FR 2956856. The twisting of such an adaptive-twist blade with reduced torsional rigidity is, on the one hand, obtained by construction, and, on the other hand, modified while the aircraft is in flight.
Nevertheless, according to a second embodiment of the invention, this blade with reduced torsional rigidity may be adapted to a traditional blade—namely, a blade that obeys a twist law, and whose twist is obtained by construction and is permanently set.
Regardless of the embodiment of the invention, the outer covering of this blade includes an extrados skin and an intrados skin, which may jointly constitute either one single skin or two separate skins. This outer covering defines a cavity that is located between the extrados skin and the intrados skin.